JP2001132405A - Device for reducing thermal stress in turbine aerofoil - Google Patents
Device for reducing thermal stress in turbine aerofoilInfo
- Publication number
- JP2001132405A JP2001132405A JP2000213617A JP2000213617A JP2001132405A JP 2001132405 A JP2001132405 A JP 2001132405A JP 2000213617 A JP2000213617 A JP 2000213617A JP 2000213617 A JP2000213617 A JP 2000213617A JP 2001132405 A JP2001132405 A JP 2001132405A
- Authority
- JP
- Japan
- Prior art keywords
- spar
- turbine airfoil
- spar structure
- turbine
- cavity
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/80—Platforms for stationary or moving blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2300/00—Materials; Properties thereof
- F05D2300/20—Oxide or non-oxide ceramics
- F05D2300/21—Oxide ceramics
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2300/00—Materials; Properties thereof
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/603—Composites; e.g. fibre-reinforced
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Architecture (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【０００１】[0001]
【発明の属する技術分野】本発明は概してエーロフォイ
ルに関し、より具体的には分離された翼桁を備えるター
ビンエーロフォイルに関する。FIELD OF THE INVENTION The present invention relates generally to airfoils, and more particularly to turbine airfoils with separated spar.
【０００２】[0002]
【従来の技術】タービンエーロフォイルは翼先端,翼長,
及び翼付根を含む。一般的に言って、冷却装置が加圧空
気をエーロフォイル翼の内部に供給する。その冷却装置
によって作り出される内部圧力はエーロフォイル翼の外
板にバルーニング応力を発生させる。その内部圧力がエ
ーロフォイル翼を損傷するのを防止するために、一般的
に外板はエーロフォイルの長手方向に延びる剛性のある
翼桁で支持される。2. Description of the Related Art Turbine airfoils have a blade tip, a blade length,
And wing roots. Generally speaking, a cooling device supplies pressurized air to the interior of the airfoil wing. The internal pressure created by the cooling device creates ballooning stress on the skin of the airfoil wing. To prevent its internal pressure from damaging the airfoil wing, the skin is typically supported by a rigid spar extending longitudinally of the airfoil.
【０００３】タービンエーロフォイルの外側表面は、作
動中には高温ガスの流れにさらされる。タービンエーロ
フォイルを冷却することで、タービンエーロフォイルの
有効寿命が延び、タービンエーロフォイルの性能が向上
する。タービンエーロフォイルの性能を向上させること
で、それを組み込んだタービンエンジンの効率と性能が
高まる。エンジンの性能が高められるのに伴って、ター
ビンエーロフォイルはより大きな空力負荷及びより高温
ガスの流れにさらされる。そのような負荷や温度に耐え
るためには、タービンエーロフォイルは複合材料を用い
て製造されればよい。そのような複合材料は負荷や高温
に耐えることができるけれども、そのような材料は普通
他の従来の材料ほど高い温度勾配に耐性がない。[0003] The outer surface of the turbine airfoil is exposed to a stream of hot gas during operation. Cooling the turbine airfoil increases the useful life of the turbine airfoil and improves the performance of the turbine airfoil. Improving the performance of turbine airfoils increases the efficiency and performance of turbine engines incorporating them. As engine performance increases, the turbine airfoil is exposed to higher aerodynamic loads and hotter gas flows. To withstand such loads and temperatures, the turbine airfoil may be manufactured using a composite material. Although such composites can withstand loads and high temperatures, such materials are generally not as resistant to high temperature gradients as other conventional materials.
【０００４】作動中には、タービンエーロフォイルは加
圧された冷却装置で内部が冷却される。従って、連続し
た翼桁はタービンエーロフォイルの外部外板表面の運転
温度より実質的に低い温度で作動している。連続した翼
桁と外部外板表面との間の温度勾配は、連続した翼桁と
外部外板表面の両方に反対の熱歪をつくり出す。温度勾
配によりつくり出される熱歪の違いにより、より低温で
作動する連続した翼桁に引張りを生じ、外部外板表面に
圧縮を生じる。セラミックスのような複合材料は高温で
は高い弾性率と低い延性を持ち、そして熱応力により連
続した翼桁に亀裂を生じ、結果としてタービンエーロフ
ォイルの破損にまで至る可能性がある。[0004] In operation, the turbine airfoil is internally cooled by a pressurized cooling device. Thus, the continuous spar operates at a temperature substantially lower than the operating temperature of the outer skin surface of the turbine airfoil. The temperature gradient between the continuous spar and the outer skin surface creates opposite thermal strains on both the continuous spar and the outer skin surface. The difference in thermal strain created by the temperature gradient causes tension in the continuous spar operating at lower temperatures and compression on the outer skin surface. Composite materials such as ceramics have high modulus and low ductility at high temperatures, and thermal stress can cause cracks in the continuous spar, resulting in failure of the turbine airfoil.
【０００５】[0005]
【発明の開示】例示的な実施形態においては、タービン
エーロフォイルは、タービンエーロフォイル内の熱応力
を減ずる分割翼桁構造を含む。タービンエーロフォイル
は、翼先端,翼付根,及び翼先端と翼付根の間に延在する
翼スパンを含む。翼スパンは、翼スパンを覆うように延
在する外板被覆と、翼スパンの長さより短い長さを有し
翼付根と翼先端の間に位置する少なくとも１つの翼桁構
造を含む。翼桁構造は第１側面と第２側面を有する少な
くとも第１翼桁を含む複数の翼桁を含む。SUMMARY OF THE INVENTION In an exemplary embodiment, a turbine airfoil includes a split spar structure that reduces thermal stresses in the turbine airfoil. The turbine airfoil includes a wing tip, a wing root, and a wing span extending between the wing tip and the wing root. The wing span includes a skin coating extending over the wing span and at least one spar structure having a length less than the length of the wing span and located between the root of the wing and the tip of the wing. The spar structure includes a plurality of spars including at least a first spar having a first side and a second side.
【０００６】作動中には、タービンエーロフォイルは、
外部外板被覆表面が分割翼桁構造の温度より高温で作動
するように内部冷却され、分離された翼桁と外部外板被
覆表面との間で温度勾配が生ずる。エーロフォイルは分
割翼桁構造を用いているので、タービンエーロフォイル
の外板表面は分割翼桁構造の間で熱膨張することがで
き、この熱膨張により外部外板表面がより高温で作動し
ていることによる熱応力の発生を防止する。従って、外
板被覆と分割翼桁構造は、従来のタービンエーロフォイ
ルでは損傷に至る可能性があった熱歪にさらされること
がなく、強度が低く延性も低い材料を用いて信頼性があ
り費用効果のよい翼桁を含むタービンエーロフォイルを
製造することができる。In operation, the turbine airfoil is
The outer skin coating surface is internally cooled to operate above the temperature of the split spar structure, creating a temperature gradient between the separated spar and the outer skin coating surface. Since the airfoil uses a split spar structure, the skin surface of the turbine airfoil can thermally expand between the split spar structures, and this thermal expansion causes the outer skin surface to operate at higher temperatures. To prevent the occurrence of thermal stress. Therefore, the skin cladding and split spar structure are not exposed to thermal strain, which could lead to damage in conventional turbine airfoils, and are reliable and cost effective using materials with low strength and low ductility. A turbine airfoil including an effective spar can be manufactured.
【０００７】[0007]
【発明の実施の形態】図１は分割翼桁構造１１を含むタ
ービンエーロフォイル１０の斜視図である。タービンエ
ーロフォイル１０は、翼付根１２、翼先端１４及び翼付
根１２と翼先端１４との間に延在する翼スパン１６を含
む。翼スパン１６は長さ１８を有し、翼付根１２から翼
先端１４まで翼スパン１６を覆うように延在する外板被
覆２０を含む。外板被覆２０は外部外板表面２１と内部
外板表面（図１には示されていない）を含む。翼長１８
は翼付根１２と翼先端１４とにかけて線２２に沿って延
びる。ある実施形態では、長さ１８は約２.０インチで
ある。タービンエーロフォイル１０はタービンエーロフ
ォイル１０を関連のタービンエンジン（図示されていな
い）に固定するように構成される取付け形状特徴２４か
ら延びている。ある実施形態では、取付け形状特徴２４
はダブテールキーである。FIG. 1 is a perspective view of a turbine airfoil 10 including a split spar structure 11. The turbine airfoil 10 includes a wing root 12, a wing tip 14, and a wing span 16 extending between the wing root 12 and the wing tip 14. Wing span 16 has a length 18 and includes a skin coating 20 that extends over wing span 16 from wing root 12 to wing tip 14. The skin coating 20 includes an outer skin surface 21 and an inner skin surface (not shown in FIG. 1). Wingspan 18
Extends along line 22 between wing root 12 and wing tip 14. In one embodiment, length 18 is about 2.0 inches. The turbine airfoil 10 extends from a mounting feature 24 configured to secure the turbine airfoil 10 to an associated turbine engine (not shown). In some embodiments, the mounting features 24
Is a dovetail key.
【０００８】図２は分割翼桁構造１１を含むタービンエ
ーロフォイル１０の部分斜視図である。タービンエーロ
フォイル１０は負圧側面５２と正圧側面５４とを含む。
正圧側面５４は負圧側面５２よりも大きい湾曲を有す
る。タービンエーロフォイル１０が空気の流れにさらさ
れるとき、正圧側面５４の大きい湾曲により、タービン
エーロフォイル１０の負圧側面５２近くに低圧の領域を
生じ、タービンエーロフォイル１０の正圧側面５４近く
に高圧の領域を生じる。FIG. 2 is a partial perspective view of the turbine airfoil 10 including the split spar structure 11. Turbine airfoil 10 includes a suction side 52 and a pressure side 54.
The pressure side 54 has a greater curvature than the suction side 52. When the turbine airfoil 10 is exposed to an airflow, the large curvature of the pressure side 54 creates a region of low pressure near the suction side 52 of the turbine airfoil 10 and near the pressure side 54 of the turbine airfoil 10. This produces high pressure areas.
【０００９】タービンエーロフォイル１０は、翼桁構造
１１が外板被覆２０と一体的に接合されて、外板被覆２
０から延びるように作られる。従って、タービンエーロ
フォイル１０の負圧側面５２は外部外板表面２１及内部
外板表面５６とを含み、タービンエーロフォイル１０の
正圧側面５４は外部外板表面２１及び内部外板表面６０
とを含む。正圧側面５４及び負圧側面５２は翼桁構造１
１に接合されてタービンエーロフォイル前縁６４及び後
縁６６を画定する。前縁６４は滑らかで負圧側面５２と
正圧側面５４との間に延在する。前縁６４は後縁６６の
幅７２より大きい幅７０を有する。The turbine airfoil 10 has a spar structure 11 integrally joined with a skin coating 20 to form a skin coating 2.
Made to extend from zero. Accordingly, the suction side 52 of the turbine airfoil 10 includes the outer skin surface 21 and the inner skin surface 56, and the pressure side 54 of the turbine airfoil 10 includes the outer skin surface 21 and the inner skin surface 60.
And The pressure side 54 and the suction side 52 have a spar structure 1
1 to define a turbine airfoil leading edge 64 and a trailing edge 66. The leading edge 64 is smooth and extends between the suction side 52 and the pressure side 54. The leading edge 64 has a width 70 that is greater than the width 72 of the trailing edge 66.
【００１０】分割翼桁構造１１は第１翼桁８０と、この
第１翼桁８０と後縁６６との間に位置する第２翼桁８２
とを含む。第１翼桁８０は第１側面８４及び第２側面８
６を有する。第１空洞８８は前縁６４と第１翼桁の第１
側面８４との間に形成される。第１翼桁８０は負圧側の
内部外板表面５６から正圧側の内部外板表面６０まで幅
９０の間を延びる。また、第１翼桁８０は翼桁構造１１
の第１側面９３から線２２に実質的に平行な方向に翼桁
構造１１の第２側面（図示されていない）に至る長さ９
２を有する。ある実施形態において、幅９０は約０.５
インチで、長さ９２は約０.２５インチである。The split spar structure 11 includes a first spar 80 and a second spar 82 located between the first spar 80 and the trailing edge 66.
And The first spar 80 includes a first side surface 84 and a second side surface 8.
6. The first cavity 88 includes the leading edge 64 and the first spar
It is formed between the side surface 84. The first spar 80 extends between the widths 90 from the suction side inner skin surface 56 to the pressure side inner skin surface 60. The first spar 80 is a spar structure 11.
From the first side 93 of the spar structure 11 to a second side (not shown) in a direction substantially parallel to the line 22.
2 In some embodiments, width 90 is about 0.5.
In inches, length 92 is about 0.25 inches.
【００１１】第２翼桁８２は第１側面９４及び第２側面
９６を有する。第２空洞９８は第１翼桁の第２側面８
６、第２翼桁の第１側面９４、正圧側の内部外板表面６
０及び負圧側の内部外板表面５６との間に形成される。
負圧側の内部外板表面５６及び正圧側の内部外板表面６
０は接合されて後縁壁１００を形成する。負圧側の外部
外板表面２１及び正圧側の外部外板表面２１は後縁壁１
００から延び、後縁６６を形成する。第３空洞１１０は
負圧側の内部外板表面５６、正圧側の内部外板表面６
０、後縁壁１００及び第２翼桁の第２側面９６の間に形
成される。第２空洞９８は第１空洞８８と第３空洞１１
０との間に位置する。The second spar 82 has a first side 94 and a second side 96. The second cavity 98 is formed on the second side surface 8 of the first spar.
6, the first side surface 94 of the second spar, the inner skin surface 6 on the pressure side
It is formed between the inner skin surface 56 on the zero and suction sides.
Suction side inner skin surface 56 and positive pressure side inner skin surface 6
0 are joined to form trailing edge wall 100. The outer skin surface 21 on the negative pressure side and the outer skin surface 21 on the positive pressure side are
00 and form a trailing edge 66. The third cavity 110 has an inner skin surface 56 on the suction side and an inner skin surface 6 on the pressure side.
0, formed between the trailing edge wall 100 and the second side 96 of the second spar. The second cavity 98 includes the first cavity 88 and the third cavity 11.
0.
【００１２】第２翼桁８２は翼桁構造１１の第１側面９
３から翼桁構造１１の第２側面に至る長さ１１２を有す
る。また、第２翼桁８２は負圧側の内部外板表面５６か
ら正圧側の内部外板表面６０に至る幅１１４を有する。
ある実施形態において、長さ１１２は第１翼桁８０の長
さ９２と実質的に同じである。また、第２翼桁８２の長
さ１１２を第１翼桁８０の長さ９２と異ならせてもよ
い。別の実施形態においては、第１翼桁８０は第２翼桁
８２と方向２２においてオフセットされる。さらに別の
実施形態においては、長さ１１２は約０.３インチで、
幅１１４は約０.３インチであり、そして第１翼桁８０
は方向２２において約０.１インチ第２翼桁８２からオ
フセットされる。The second spar 82 is a first side surface 9 of the spar structure 11.
It has a length 112 from 3 to the second side of the spar structure 11. The second spar 82 also has a width 114 from the suction side internal skin surface 56 to the pressure side internal skin surface 60.
In some embodiments, length 112 is substantially the same as length 92 of first spar 80. Further, the length 112 of the second spar 82 may be different from the length 92 of the first spar 80. In another embodiment, first spar 80 is offset in direction 22 with second spar 82. In yet another embodiment, length 112 is about 0.3 inches,
The width 114 is about 0.3 inches and the first spar 80
Is offset from the second spar 82 by about 0.1 inch in the direction 22.
【００１３】作動中には、外部外板表面２１は高温ガス
の流れにさらされる。タービンエーロフォイル１０を冷
却するために、冷却装置（図示されていない）が加圧空
気の流れをタービンエーロフォイル１０の内部に供給す
る。冷却装置により供給される加圧空気の流れによっ
て、翼桁構造１１は、外部外板表面２１、正圧側の内部
外板表面６０及び負圧側の内部外板表面５６を含む外板
被覆２０よりかなり冷たい温度で作動する。従って、温
度勾配が外板被覆２０と翼桁構造１１との間に発生す
る。In operation, the outer skin surface 21 is exposed to a flow of hot gas. To cool the turbine airfoil 10, a cooling device (not shown) supplies a flow of pressurized air to the interior of the turbine airfoil 10. Due to the flow of pressurized air provided by the cooling device, the spar structure 11 is significantly more than the skin coating 20 including the outer skin surface 21, the inner skin surface 60 on the pressure side and the inner skin surface 56 on the suction side. Operates at cold temperatures. Therefore, a temperature gradient is generated between the skin coating 20 and the spar structure 11.
【００１４】翼桁構造の翼桁８０及び８２はそれぞれ長
さ９２及び１１２を有し、それらが翼桁構造１１に熱歪
を生じさせることなく、正圧側面５４及び負圧側面５２
が熱膨張することを可能にする。その結果、翼桁構造１
１は強度が低く延性も低い材料で作ることが可能であ
る。ある実施形態において、翼桁構造１１はＳｉＣ−Ｓ
ｉＣセラミックマトリックス複合材料で造られる。ま
た、翼桁構造１１はモノリシックセラミック材料から造
られる。The spar structure spar 80 and 82 have lengths 92 and 112, respectively, so that they do not cause thermal strain in the spar structure 11 and the pressure side 54 and suction side 52
Allow for thermal expansion. As a result, the spar structure 1
No. 1 can be made of a material having low strength and low ductility. In one embodiment, the spar structure 11 is made of SiC-S
Made of iC ceramic matrix composite. Also, the spar structure 11 is made from a monolithic ceramic material.
【００１５】また、タービンエーロフォイル１０には追
加の翼桁構造１２０を設けることができる。翼桁構造１
２０は翼桁構造１１と実質的に同様に構成され、第１翼
桁１２２及び第２翼桁１２４とを含む。翼桁構造１２０
は翼桁構造１１と翼先端１４の間に配置され、翼桁１２
２及び１２４はそれぞれ翼桁構造１１から距離１２６及
び１２８をおいて位置される。ある実施形態において
は、翼桁構造１２０は翼桁構造１１から約０.１インチ
のところに配置される。別の実施形態では、第１翼桁１
２２は第１翼桁８０から方向１２９の方にオフセットさ
れ、第２翼桁１２４は第２翼桁８２から方向１２９の方
にオフセットされる。ある実施形態では、翼桁１２２及
び１２４はそれぞれ翼桁８０及び８２から方向１２９の
方に約０.１インチにオフセットされる。Further, the turbine airfoil 10 may be provided with an additional spar structure 120. Spar structure 1
Reference numeral 20 is substantially similar to the spar structure 11 and includes a first spar 122 and a second spar 124. Spar structure 120
Is located between the spar structure 11 and the wing tip 14 and the spar 12
2 and 124 are located at distances 126 and 128 from the spar structure 11, respectively. In one embodiment, spars structure 120 is located about 0.1 inches from spars structure 11. In another embodiment, the first spar 1
22 is offset from the first spar 80 in the direction 129, and the second spar 124 is offset from the second spar 82 in the direction 129. In one embodiment, spars 122 and 124 are offset from spars 80 and 82 by about 0.1 inch in direction 129, respectively.
【００１６】図３は分割翼桁構造１３２を含むタービン
エーロフォイル１３０の部分斜視図である。ある実施形
態において、タービンエーロフォイル１３０はフレーム
ストラットである。タービンエーロフォイル１３０は翼
先端（図示されていない）、翼付根（図示されていな
い）を含み、そして翼付根と翼先端の間に延在する翼ス
パン１３６を有する。さらに、タービンエーロフォイル
１３０は第１側面１４０と第２側面１４２とを含む。タ
ービンエーロフォイル１３０は翼スパン１３６を覆って
延在する外部外板被覆表面１４４を含む。第１側面１４
０は外部外板被覆表面１４４と内部外板表面１４６とを
含む。タービンエーロフォイル１３０の第２側面１４２
は外部外板表面１４４と内部外板表面１４８とを含む。
第１側面１４０及び第２側面１４２は翼桁構造１３２に
接合されて、タービンエーロフォイルの前縁１５０を画
定する。前縁１５０は滑らかで、第１側面１４０と第２
側面１４２の間に延在する。外部外板表面１４４は前縁
１５０から後縁１５２に延在する。タービンエーロフォ
イルの第１側面１４０は、第２側面１４２に延在する湾
曲と実質的に同一である、前縁１５０から後縁１５２ま
で延在する湾曲を有する。ある実施形態においては、タ
ービンエーロフォイル１３０は対称形のエーロフォイル
である。FIG. 3 is a partial perspective view of the turbine airfoil 130 including the split spar structure 132. In some embodiments, turbine airfoil 130 is a frame strut. The turbine airfoil 130 includes a wing tip (not shown), a wing root (not shown), and has a wing span 136 extending between the wing root and the wing tip. Further, turbine airfoil 130 includes a first side 140 and a second side 142. Turbine airfoil 130 includes an outer skin coating surface 144 that extends over blade span 136. First side 14
0 includes an outer skin coating surface 144 and an inner skin surface 146. Second side 142 of turbine airfoil 130
Includes an outer skin surface 144 and an inner skin surface 148.
First side 140 and second side 142 are joined to spar structure 132 to define a leading edge 150 of the turbine airfoil. The leading edge 150 is smooth, with the first side 140 and the second
It extends between the side surfaces 142. Outer skin surface 144 extends from leading edge 150 to trailing edge 152. The first side 140 of the turbine airfoil has a curvature extending from a leading edge 150 to a trailing edge 152 that is substantially identical to the curvature extending to the second side 142. In some embodiments, turbine airfoil 130 is a symmetric airfoil.
【００１７】分割翼桁構造１３２は第１翼桁１６０及び
この第１翼桁１６０と後縁１５２との間に位置する第２
翼桁１６２とを含む。第１翼桁１６０は第１側面１６４
及び第２側面１６６を有する。第１空洞１６８は前縁１
５０と第１翼桁の第１側面１６４との間に形成される。
第１翼桁１６０は第１側の内部外板表面１４６から第２
側の内部外板表面１４８にかけて幅１７０の間を延び
る。また、第１翼桁１６０は翼桁構造１３２の第１側面
１７３から翼桁構造１３２の第２側面（図示されていな
い）に至る長さ１７２を有する。The split spars structure 132 includes a first spars 160 and a second spars located between the first spars 160 and the trailing edge 152.
Spar 162. The first spar 160 is the first side surface 164
And a second side surface 166. The first cavity 168 is the leading edge 1
It is formed between 50 and the first side surface 164 of the first spar.
The first spar 160 is second from the inner skin surface 146 on the first side.
Extending between the widths 170 across the inner skin surface 148 on the side. The first spar 160 also has a length 172 from the first side 173 of the spar structure 132 to the second side (not shown) of the spar structure 132.
【００１８】第２翼桁１６２は第１側面１８０と第２側
面１８２を有する。第２空洞１８４は第１翼桁の第２側
面１６６、第２翼桁の第１側面１８０、第１側の内部外
板表面１４６及び第２側の内部外板表面１４８の間に形
成される。第３空洞１８５は第２翼桁の第２側面１８
２、第１側の内部外板表面１４６、後縁１５２及び第２
側の内部外板表面１４８の間に形成される。第２翼桁１
６２は翼桁構造１３２の第１側面１７３から翼桁構造１
３２の第２側面に至る長さ１８８を有する。また第２翼
桁１６２は第２側の内部外板表面１４８から第１側の内
部外板表面１４６に至る幅１９０を有する。The second spar 162 has a first side 180 and a second side 182. The second cavity 184 is formed between the second side 166 of the first spar, the first side 180 of the second spar, the inner skin surface 146 on the first side, and the inner skin surface 148 on the second side. . The third cavity 185 is formed on the second side surface 18 of the second spar.
2, the inner skin surface 146 on the first side, the trailing edge 152 and the second
Formed between the side inner skin surfaces 148. 2nd spar 1
62 denotes a spar structure 1 from the first side surface 173 of the spar structure 132.
It has a length 188 leading to 32 second sides. The second spar 162 also has a width 190 from the second side inner skin surface 148 to the first side inner skin surface 146.
【００１９】図４は分割翼桁構造２０２を含む高圧ベー
ン２００の斜視図である。ベーン２００はベーン付根２
０４、ベーン先端２０６及びベーン付根２０４とベーン
先端２０６との間に延在するベーンスパン２０８とを含
む。ベーンスパン２０８は長さ２１０を有し、ベーンス
パン２０８を覆うようにベーン付根２０４からベーン先
端２０６まで延在する外板被覆２１２を含む。外板被覆
２１２は外部外板表面２１４及び内部外板表面（図示さ
れていない）を含む。高圧ベーン２００はベーン２００
を固定するよう構成された取付け形状特徴２２０から延
在する。FIG. 4 is a perspective view of a high pressure vane 200 including a split spar structure 202. Vane 200 is the base 2 of the vane
04, a vane tip 206 and a vane pan 208 extending between the vane root 204 and the vane tip 206. Vane span 208 has a length 210 and includes skin cladding 212 extending from vane root 204 to vane tip 206 to cover vane pan 208. Skin coating 212 includes an outer skin surface 214 and an inner skin surface (not shown). High pressure vane 200 is vane 200
Extend from a mounting feature 220 configured to secure the
【００２０】分割翼桁構造２０２は第１翼桁２２２と第
２翼桁２２４とを含む。第１翼桁２２２は第１空洞２３
０と第２空洞２２８の間に配置される。第２翼桁２２４
は空洞２２８と第３空洞２２６との間に配置される。The split spar structure 202 includes a first spar 222 and a second spar 224. The first spar 222 is the first cavity 23
0 and the second cavity 228. Second spar 224
Is located between cavity 228 and third cavity 226.
【００２１】図５は分割翼桁構造２５２を含むストラッ
ト前縁エクステンション２５０の斜視図である。ストラ
ット前縁エクステンション２５０は、第１端２５４、第
２端（図示されていない）及び第１端２５４と第２端の
間に延在するエクステンションスパン２５６を有する。
外板被覆２５８はエクステンション２５０を覆うように
第１端２５４から第２端まで延在し、前縁２６０と後縁
２６２とを画定する。後縁２６２はストラット前縁エク
ステンション２５０をストラット（図示されていない）
に固定するように構成された取付け形状特徴２６４まで
延びている。ある実施形態においては、取り付け形状特
徴２６４はダブテールキーである。FIG. 5 is a perspective view of a strut leading edge extension 250 that includes a split spar structure 252. The strut leading edge extension 250 has a first end 254, a second end (not shown), and an extension span 256 extending between the first end 254 and the second end.
The skin cladding 258 extends from the first end 254 to the second end over the extension 250 and defines a leading edge 260 and a trailing edge 262. The trailing edge 262 connects the strut leading edge extension 250 to the strut (not shown).
Extend to a mounting feature 264 that is configured to be secured to the mounting feature. In some embodiments, mounting feature 264 is a dovetail key.
【００２２】分割翼桁構造２５２は第１翼桁部分２７０
を含む。第１翼桁部分２７０は第１側面２７２、第２側
面２７３及び長さ２７４を有する。第１翼桁部分２７０
は、ストラット前縁エクステンションのスパン２５６に
沿って分割距離２７６で分割されており、第2の部分２
７８を有する。第１側面２７２は第１空洞２７９の境界
となり、そして第２側面２７３は第２空洞２８０の境界
となる。第１翼桁２７０は外板被覆２５８と一体に形成
され、ストラット前縁エクステンション２５０の第１側
面２８２からストラット前縁エクステンション２５０の
第２側面２８４まで延びる。従って、分割翼桁構造２５
２の全体の翼桁の長さは、第２の部分２７８の長さと第
１の部分２７０の長さ２７４との合計に等しく、そして
この全体の翼桁の長さはスパン２５６より小さい。The split spar structure 252 includes a first spar portion 270.
including. First spar portion 270 has a first side 272, a second side 273, and a length 274. First spar part 270
Are separated by a split distance 276 along the span 256 of the strut leading edge extension, and the second part 2
78. The first side 272 bounds the first cavity 279 and the second side 273 bounds the second cavity 280. The first spar 270 is formed integrally with the skin cladding 258 and extends from a first side 282 of the strut leading edge extension 250 to a second side 284 of the strut leading edge extension 250. Therefore, the split spar structure 25
The length of the entire spars is equal to the sum of the length of the second portion 278 and the length 274 of the first portion 270, and the length of the entire spars is less than the span 256.
【００２３】上記のタービンエーロフォイルは経済的で
信頼性のある分割翼桁構造を含む。タービンエーロフォ
イルは、タービンエーロフォイルの翼全長より短い全長
を有し、エーロフォイル外板を冷却装置によって発生す
る内部圧力に対して支持するために複数の翼桁を含む少
なくとも１つの翼桁構造を含む。さらに、その翼桁構造
はタービンエーロフォイルの外部外板表面が熱膨張する
のを可能にする。そのような膨張はタービンエーロフォ
イル内に熱歪が生じるのを防止し、翼桁構造を強度が低
く延性も低い材料で造ることを可能にする。従って、経
済的で精度のよいエーロフォイル翼桁構造が得られる。The above-described turbine airfoil includes an economical and reliable split spar structure. The turbine airfoil has an overall length that is less than the turbine airfoil wing length, and includes at least one spar structure including a plurality of spars for supporting the airfoil skin against internal pressure generated by the cooling device. Including. In addition, the spar structure allows the outer skin surface of the turbine airfoil to thermally expand. Such expansion prevents thermal strain from occurring in the turbine airfoil and allows the spar structure to be made of a low strength, low ductility material. Therefore, an economical and accurate airfoil spar structure can be obtained.
【００２４】本発明をこれまで様々な特定の実施形態に
ついて説明してきたが、当業者は本発明をその請求項の
精神と範囲内の変更形態でもって実施し得ることが理解
できるであろう。While the present invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the appended claims.
【図１】 分割翼桁構造を含むタービンエーロフォイル
の斜視図。FIG. 1 is a perspective view of a turbine airfoil including a split spar structure.
【図２】 図１に示される線２−２に沿うタービンエー
ロフォイルの横断面図。FIG. 2 is a cross-sectional view of the turbine airfoil taken along line 2-2 shown in FIG.
【図３】 分割翼桁構造を含むタービンエーロフォイル
の別の実施形態の横断面図。FIG. 3 is a cross-sectional view of another embodiment of a turbine airfoil including a split spar structure.
【図４】 分割翼桁構造を含む高圧ベーンの斜視図。FIG. 4 is a perspective view of a high-pressure vane including a split-spar structure.
【図５】 分割翼桁構造を含むストラット前縁エクステ
ンションの斜視図。FIG. 5 is a perspective view of a strut leading edge extension including a split spar structure.
───────────────────────────────────────────────────── フロントページの続き (72)発明者 マーク・ユージーン・ノー アメリカ合衆国、オハイオ州、モロウ、ウ ェッジウッド・ドライブ、5895番 (72)発明者 ダグラス・メルトン・カーパー アメリカ合衆国、オハイオ州、ブルー・ア ッシュ、ナンバー104、ハント・ロード、 4894番 ──────────────────────────────────────────────────の Continued on front page (72) Inventor Mark Eugene No United States, Ohio, Morrow, Wedgewood Drive, No. 5895 (72) Inventor Douglas Melton Carper Blue Ash, Ohio, United States of America , Number 104, hunt road, number 4894
Claims (19)
６）と、 前記翼スパンの長さより短い長さ（９２）を有し、前記
翼付根と前記翼先端との間に位置し、前記タービンエー
ロフォイル第１側面から前記タービンエーロフォイル第
２側面に至る幅（９０）をもつ第１の翼桁（８０）を含
む複数の翼桁（８０，８２）を有する少なくとも１つの
翼桁構造（１１）とを含んでなるタービンエーロフォイ
ル（１０）。A wing root (12); a wing tip (14); a first side surface (54); a second side surface (52) laterally opposed to said first side surface; The wing span (1) extending between the wing tip
6) having a length (92) shorter than the length of the blade span, located between the root of the blade and the tip of the blade, from the first side of the turbine airfoil to the second side of the turbine airfoil. At least one spar structure (11) having a plurality of spars (80, 82), including a first spar (80, 82) having a width (90) leading to the turbine airfoil (10).
ロフォイルの熱応力を減ずるように構成される請求項１
に記載のタービンエーロフォイル（１０）。2. The spar structure (11) is configured to reduce thermal stresses in the turbine airfoil.
A turbine airfoil (10) according to claim 1.
びる外板被覆（２０）を含み、前記タービンエーロフォ
イル第１側面（５４）を前記第２側面（５２）に接合し
て後縁（６６）まで延びる前縁（６４）を画定し、前記
前縁が前記後縁と軸方向に相対して位置する請求項２に
記載のタービンエーロフォイル（１０）。3. A turbine engine comprising a skin coating (20) extending over said blade span (16) and joining said turbine airfoil first side (54) to said second side (52) for trailing edge. A turbine airfoil (10) in accordance with Claim 2 defining a leading edge (64) extending to (66), said leading edge being axially opposed to said trailing edge.
面（８６）とを備え、前記第１側面が第１空洞（８８）
の境界となり、前記第１翼桁の第２側面が第２空洞（９
８）の境界になる請求項３に記載のタービンエーロフォ
イル（１０）。4. The first spar has a first side (84) and a second side (86), the first side being a first cavity (88).
And the second side surface of the first spar is in the second cavity (9).
The turbine airfoil (10) according to claim 3, which is bounded by (8).
（９４）と第２側面（９６）を備える第２翼桁（８２）
をさらに含む請求項４に記載のタービンエーロフォイル
（１０）。5. A second spar (82) wherein said plurality of spar (80, 82) comprises a first side (94) and a second side (96).
The turbine airfoil (10) according to claim 4, further comprising:
２空洞（９８）の境界となり、そして前記第２翼桁の第
２側面（９６）が第３空洞（１１０）の境界となる請求
項５に記載のタービンエーロフォイル（１０）。6. The second spar first side (94) borders the second cavity (98) and the second spar second side (96) is defined by a third cavity (110). The turbine airfoil (10) according to claim 5, which is a boundary.
低い材料からなる請求項４に記載のタービンエーロフォ
イル（１０）。7. The turbine airfoil (10) according to claim 4, wherein said spar structure (11) is made of a material having low strength and low ductility.
ックス複合材料からなる請求項４に記載のタービンエー
ロフォイル（１０）。8. A turbine airfoil (10) according to claim 4, wherein said spar structure (11) comprises a ceramic matrix composite.
ミック材料からなる請求項４に記載のタービンエーロフ
ォイル（１０）。9. The turbine airfoil (10) according to claim 4, wherein said spar structure (11) comprises a monolithic ceramic material.
０）を有し、前記第２翼桁（８２）が前記タービンエー
ロフォイル第１側面（５４）から前記タービンエーロフ
ォイル第２側面（５２）に至る第２の幅（１１４）を有
する請求項５に記載のタービンエーロフォイル（１
０）。10. The first spar (80) has a first width (9).
The second spar (82) has a second width (114) from the turbine airfoil first side (54) to the turbine airfoil second side (52). Turbine airfoil (1)
0).
第２翼桁の第２の幅（１１４）は、前記タービンエーロ
フォイル第２側面（５２）が前記第１側面（５４）より
大きい湾曲を有するように構成される請求項１０に記載
のタービンエーロフォイル（１０）。11. The first width (90) of the first spar and the second width (114) of the second spar are such that the second side surface (52) of the turbine airfoil is the first side surface (52). 54) The turbine airfoil (10) of claim 10, wherein the turbine airfoil (10) is configured to have a greater curvature.
第２翼桁の第２の幅（１１４）は、前記タービンエーロ
フォイル第２側面（５２）が前記タービンエーロフォイ
ル第１側面（５４）の湾曲と同一の湾曲を有するように
構成される請求項１０に記載のタービンエーロフォイル
（１０）。12. The first width (90) of the first spar and the second width (114) of the second spar are such that the second side surface (52) of the turbine airfoil is the same as that of the turbine airfoil. The turbine airfoil (10) according to claim 10, wherein the turbine airfoil (10) is configured to have the same curvature as one side (54).
を有し、翼付根（１２）、翼先端（１４）及び該翼先端
と該翼付根との間に延在する翼スパン（１６）とを含む
タービンエーロフォイル（１０）において、タービンエ
ーロフォイル内の熱応力を減ずるように構成された翼桁
構造（１１）であって、 前記タービンエーロフォイル第１側面と第２側面の間に
延びている少なくとも第１翼桁（８０）を含み、前記翼
スパンの長さより短い長さ（９２）をもつ複数の翼桁
（８０，８２）を含んでなる翼桁構造（１１）。13. A blade having a first side (54) and a second side (52) and extending between the blade root (12), the blade tip (14) and the blade tip and the blade root. A spar structure (11) configured to reduce thermal stress in the turbine airfoil, the turbine airfoil (10) including a span (16), the turbine airfoil having a first side and a second side. A spar structure (11) comprising at least a first spar (80) extending therebetween and a plurality of spars (80, 82) having a length (92) shorter than the length of the wing span. .
（１０）を覆うように延在する外板被覆（２０）を含
み、前記外板被覆から延びている請求項１３に記載の翼
桁構造（１１）。14. The spar structure (11) according to claim 13, further comprising a skin coating (20) extending over the turbine airfoil (10) and extending from the skin coating. .
４）及び第２側面（８６）を備えており、前記第１側面
が第１空洞（８８）の境界となり、前記第２側面が第２
空洞（９８）の境界となる請求項１４に記載の翼桁構造
（１１）。15. The first spar (80) has a first side surface (8).
4) and a second side surface (86), wherein the first side surface is a boundary of the first cavity (88), and the second side surface is a second side surface (86).
15. The spar structure (11) according to claim 14, bordering the cavity (98).
が第１側面（９４）及び第２側面（９６）を備える第２
翼桁（８２）を含んでおり、前記第２翼桁の第１側面が
前記第２空洞（９８）の境界となり、前記第２翼桁の第
２側面が第３空洞（１１０）の境界となる請求項１５に
記載の翼桁構造（１１）。16. The plurality of spar beams (80, 82).
Has a first side (94) and a second side (96).
A spar (82), wherein a first side of the second spar is a boundary of the second cavity (98) and a second side of the second spar is a boundary of a third cavity (110). Spar structure (11) according to claim 15, comprising:
らなる請求項１５に記載の翼桁構造（１１）。17. The spar structure according to claim 15, wherein the spar structure is made of a material having low strength and ductility.
複合材料からなる請求項１５に記載の翼桁構造（１
１）。18. The spar structure according to claim 15, wherein said spar structure is made of a ceramic matrix composite material.
1).
材料からなる請求項１５に記載の翼桁構造（１１）。19. The spar structure according to claim 15, wherein the spar structure is made of a monolithic ceramic material.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US09/399194 | 1999-09-17 | ||
| US09/399,194 US6398501B1 (en) | 1999-09-17 | 1999-09-17 | Apparatus for reducing thermal stress in turbine airfoils |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JP2001132405A true JP2001132405A (en) | 2001-05-15 |
| JP2001132405A5 JP2001132405A5 (en) | 2007-08-16 |
Family
ID=23578539
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP2000213617A Pending JP2001132405A (en) | 1999-09-17 | 2000-07-14 | Device for reducing thermal stress in turbine aerofoil |
Country Status (4)
| Country | Link |
|---|---|
| US (1) | US6398501B1 (en) |
| EP (1) | EP1085170B1 (en) |
| JP (1) | JP2001132405A (en) |
| DE (1) | DE60032695T2 (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2010007650A (en) * | 2008-06-30 | 2010-01-14 | Mitsubishi Heavy Ind Ltd | Turbine blade and gas turbine |
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| US7435058B2 (en) * | 2005-01-18 | 2008-10-14 | Siemens Power Generation, Inc. | Ceramic matrix composite vane with chordwise stiffener |
| US8137611B2 (en) * | 2005-03-17 | 2012-03-20 | Siemens Energy, Inc. | Processing method for solid core ceramic matrix composite airfoil |
| US7785076B2 (en) * | 2005-08-30 | 2010-08-31 | Siemens Energy, Inc. | Refractory component with ceramic matrix composite skeleton |
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| US7600978B2 (en) * | 2006-07-27 | 2009-10-13 | Siemens Energy, Inc. | Hollow CMC airfoil with internal stitch |
| US8033790B2 (en) * | 2008-09-26 | 2011-10-11 | Siemens Energy, Inc. | Multiple piece turbine engine airfoil with a structural spar |
| DK2524134T3 (en) | 2010-01-14 | 2014-08-11 | Neptco Inc | Rotor blade components for a wind turbine and methods for making them |
| US10137542B2 (en) | 2010-01-14 | 2018-11-27 | Senvion Gmbh | Wind turbine rotor blade components and machine for making same |
| US20110206522A1 (en) * | 2010-02-24 | 2011-08-25 | Ioannis Alvanos | Rotating airfoil fabrication utilizing cmc |
| WO2014133721A1 (en) * | 2013-02-27 | 2014-09-04 | United Technologies Corporation | Gas turbine engine thin wall composite vane airfoil |
| EP3059390B1 (en) * | 2015-02-18 | 2020-03-04 | Rolls-Royce Corporation | Vane assembly for a gas turbine engine, airfoil and method of making |
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| US11773723B2 (en) * | 2019-11-15 | 2023-10-03 | Rtx Corporation | Airfoil rib with thermal conductance element |
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Also Published As
| Publication number | Publication date |
|---|---|
| US6398501B1 (en) | 2002-06-04 |
| EP1085170A2 (en) | 2001-03-21 |
| EP1085170B1 (en) | 2007-01-03 |
| DE60032695D1 (en) | 2007-02-15 |
| DE60032695T2 (en) | 2007-10-31 |
| EP1085170A3 (en) | 2003-01-02 |
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